Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of nickel and cobalt-base superalloys. Nonetheless, when used to form components of the turbine, combustor and augmentor sections of a gas turbine engine, such alloys alone are often susceptible to damage by oxidation and hot corrosion attack and may not retain adequate mechanical properties. For this reason, these components are often protected by an environmental and/or thermal-insulating coating, the latter of which is termed a thermal barrier coating (TBC) system. Ceramic materials, and particularly yttria-stabilized zirconia (YSZ), are widely used as a thermal barrier coating (TBC), or topcoat, of TBC systems used on gas turbine engine components. These particular materials are widely employed because they can be readily deposited by plasma spray, flame spray and vapor deposition techniques. A commonly used type of TBC is a coating based on zirconia stabilized with yttria, for example about 93 wt. % zirconia stabilized with about 7 wt. % yttria. This general type of TBC has been reported in such United States patents as U.S. Pat. No. 4,055,705, U.S. Pat. No. 4,328,285, and U.S. Pat. No. 5,236,745, which are incorporated herein by reference. Such TBC coatings have a relatively rough surface and do not provide adequate heat energy reflection for certain applications. In addition, application of certain TBC coatings requires use of apparatus having a controlled atmosphere or vacuum. Accordingly, such coatings and methods cannot be effectively utilized in field repairs.
To be effective, TBC systems must have low thermal conductivity, strongly adhere to the component, and remain adherent throughout many heating and cooling cycles. The latter requirement is particularly demanding due to the different coefficients of thermal expansion between ceramic topcoat materials and the superalloy substrates they protect. To promote adhesion and extend the service life of a TBC system, a bond coat is often employed. Bond coats are typically in the form of overlay coatings such as MCrAlX (where M is iron, cobalt and/or nickel, and X is yttrium or another rare earth element), or diffusion aluminide coatings. During the deposition of the ceramic TBC and subsequent exposures to high temperatures, such as during engine operation, these bond coats form a tightly adherent alumina (Al2O3) layer or oxide scale that adheres the TBC to the bond coat.
The service life of a TBC system is typically limited by a spallation event brought on by thermal fatigue and stress, coating defects, mechanical damage, wear, and the like. Accordingly, a significant challenge of TBC systems has been to obtain a more adherent ceramic layer that is less susceptible to spalling when subjected to thermal cycling. Though significant advances have been made, there is the inevitable requirement to repair components whose thermal barrier coatings have spalled. Though spallation typically occurs in localized regions or patches, the conventional repair method has been to completely remove the thermal barrier coating, restore or repair the bond layer surface as necessary, and then recoat the entire component. Prior art techniques for removing TBC's include grit blasting or chemically stripping with an alkaline solution at high temperatures and pressures. However, grit blasting is a slow, labor-intensive process and erodes the surface beneath the coating. With repetitive use, the grit blasting process eventually destroys the component. The use of an alkaline solution to remove a thermal barrier coating is also less than ideal, since the process requires the use of an autoclave operating at high temperatures and pressures. Consequently, conventional repair methods are labor-intensive and expensive, and can be difficult to perform on components with complex geometries, such as airfoils and shrouds. As an alternative, U.S. Pat. No. 5,723,078 to Nagaraj et al. teaches selectively repairing a spalled region of a TBC by texturing the exposed surface of the bond coat, and then depositing a ceramic material on the textured surface by plasma spraying. While avoiding the necessity to strip the entire TBC from a component, the repair method taught by Nagaraj et al. still requires removal of the component from the engine assembly in order to deposit the ceramic material, and further requires the use of plasma spraying apparatus to effect the repair.
Moreover, existing sprayable TBC materials require some type of post drying or firing in order to be stabilized prior to high temperature use, and therefore are ineffective for in-situ field repairs. Tape materials require an autoclave to apply, and are thus not feasible for in-situ repairs. While plasma sprayed materials do not all require post-deposition heating, such materials have much rougher finishes, and cannot be applied in the field for in-situ repairs without spraying powder throughout the rest of the engine (which requires a major cleaning step prior to subsequent engine operation).
In the case of aircraft turbine engines and large power generation turbines, removing the turbine from service for repairs results in significant costs in terms of labor and downtime. For these reasons, removing components having TBCs that have suffered only localized spallation is not economically desirable. As a result, components identified as having spalled TBC are often analyzed to determine whether the spallation has occurred in a high stress area, and a judgment is then made as to the risk of damage to the turbine due to the reduced thermal protection of the component that could lead to catastrophic failure of the component. If the decision is to continue operation, the spalled component must typically be scrapped at the end of operation because of the thermal damage inflicted while operating the component without complete TBC coverage. Additionally, some newer TBCs utilize a smoothing layer over the TBC for better heat rejection and air flow. Currently, there is no known way at present to replace this smoothing layer having a very smooth finish on damaged TBC.
Accordingly, it would be desirable if a repair method were available that could be performed on localized spalled areas of TBC on turbine hardware in field and in situ, without necessitating that the component be removed from the turbine, so that downtime and scrappage are minimized.
It would also be desirable to repair a smoothing layer on a damaged TBC in a manner that restores the very smooth finish of the smoothing layer, as well as restoring the heat rejection and airflow properties of the smoothing layer.
It would further be desirable to provide an improved smoothing coating for repair of damaged TBC that is easy to apply by caulking, spackling or brushing in-situ.